The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
The core engine of a gas turbine engine typically includes a multistage axial compressor which provides compressed air to a combustor wherein it is mixed with fuel and ignited for generating hot combustion gas which flows downstream through a high pressure turbine nozzle and in turn through one or more stages of turbine rotor blades. The high pressure turbine blades are suitably joined to a rotor disk which is joined to the compressor by a corresponding drive shaft, with the turbine blades extracting energy for powering the compressor during operation in a two spool engine, a second shaft joins a fan upstream of the compressor to a low pressure turbine disposed downstream from the high pressure turbine for providing additional propulsion force for typical use in powering an aircraft in flight.
Typical turbine nozzles, such as high pressure and low pressure turbine nozzles, have fixed vane configurations and fixed nozzle throat areas there between in view of the severe temperature and high pressure loading environment in which they operate. The throat areas between adjacent nozzle vanes must be accurately maintained for maximizing performance of the engine, yet the hot thermal environment requires that the turbine nozzle be manufactured in circumferential segments for reducing thermal stress during operation. The nozzle segments therefore require suitable inter-segment sealing to reduce undesirable flow leakage, which further complicates turbine nozzle design.
Variable cycle engines are being developed for maximizing performance and efficiency over subsonic and supersonic flight conditions. Although it would be desirable to obtain variable flow through turbine nozzles by adjusting the throat areas thereof, previous attempts thereat have proved impractical in view of the severe operating environment of the nozzles. For example, it is common to provide variability in compressor stator vanes by mounting each vane on a radial spindle and collectively rotating each row of compressor vanes using an annular unison ring attached to corresponding lever arms joined to each of the spindles. In this way the entire compressor vane rotates or pivots about a radial axis, with suitable hub and tip clearances being required for permitting the vanes to pivot.
Applying the variable compressor configuration to a turbine nozzle has substantial disadvantages both in mechanical implementation as well as in aerodynamic performance. The severe temperature environment of the turbine nozzles being bathed in hot combustion gases from the combustor typically requires suitable cooling of the individual vanes, with corresponding large differential temperature gradients through the various components. A pivotable nozzle vane increases the difficulty of design, and also provides hub and tip gaps which require suitable sealing since any leakage of the combustion gas therethrough adversely affects engine performance and efficiency which negates the effectiveness of the variability being introduced.
In the related patent application identified above, an improved variable area turbine nozzle includes a plurality of stationary first vane segments which cooperate with complementary second vane segments which are pivotally joined to outer and inner bands to define respective two-piece vanes for adjusting throat area therebetween. The second vane segments are pivoted like doors relative to the corresponding first vane segments and therefore require suitable seals at their radial and axial ends to contain pressurized cooling air inside the individual vanes.
The present invention provides various forms of static seals therefor, which have additional utility wherever sealing is required between two spaced apart components subject to differential movement therebetween during operation.